This invention is in the field of control systems for spacecraft, aircraft, satellites, missiles and other movable vehicles requiring attitude control, such control being along one, two or three axes of attitude of the control vehicle.
Generally, attitude control is gyroscopically controlled in one or more directions by means of a rotating wheel of substantial mass and rotational speed to stabilize the attitude of the vehicle in a predetermined direction. The large moments of inertia thereby produced, inhibit to a large extent, rapid correction of attitude when needed since torques or forces must be developed to counteract the effect of these moments. Gimbaling systems provide capability, in the rotating wheel stabilizer, to obtain more than one degree of freedom of stabilization. However, the more degrees of freedom sought to be obtained by these gyro stabilizers the greater the complexity thereof, and a priori, the greater the unreliability of such a control system. A reliable and relatively simple system would be required considering that certain applications of these attitude control systems may be in orbital or galaxy investigation vehicles that may require many years of operative life. Most of the control system types described by the prior art are either dependent upon gravitational action of the earth or other celestial bodies or upon their geomagnetic fields, if these exist. Thus far, at least in the solar system, only two such bodies exhibit geomagnetic fields, and hence the prior art control systems are unsuitable for a vehicle whose objectives and requirements are as extensive as the one described in the instant specification.
The types of gyro systems hereinabove described are exemplified by U.S. Pat. Nos. 3,452,948; 3,582,019; 3,188,639; 3,638,883; 3,741,500; 3,171,612; 3,105,657; 3,424,401; 3,493,194 and 3,567,155.
A control system, other than those above identified by the referenced patents, is exemplified by U.S. Pat. No. 3,291,419 which utilizes the earth's magnetic field by using magnetometer devices as sensors. The obvious disadvantage of this system, as mentioned previously, is that it is limited to control systems that depend upon a celestial body which has a magnetic field, and few of such bodies are known to possess such field. Additionally, this system is also mechanically and electrically complex and possesses a high degree of unreliability.
A system used to control aircraft by thrust producers is exemplified by U.S. Pat. No. 2,943,822. This system also employs gyroscopic mechanisms, used to control several degrees of freedom, such as a vertical gyro and a directional control gyro, in addition to propulsion units to offset gravitational effects. This patent is a hybrid between gyro and thrust control and is also very complex mechanically and electrically, and thus has a high degree of built-in unreliability and is unsuitable for the purposes intended by the instant specification.
A fluid proportional thrust system is depicted in U.S. Pat. No. 3,612,442. Attitude control is achieved in the moving vehicle of this system by the use of vortex amplifying devices to modulate continuous gas flow through the several thrustors in response to electrical or fluidic control systems. One of the obvious disadvantages of this system is that constant use of fuel is required to maintain control thus limiting severely the mission period allocated to that vehicle as well as the distances traveled from the point of launch.
Another hybrid control system utilizing a spring body and thrustors is characterized by U.S. Pat. No. 3,511,452. The system illustrated by this patent depends upon revolutions of a reaction wheel in a suitable housing energized by a plurality of pulses. Jets are also provided for controlling the vehicle rotational speed. The reaction wheel is counter-rotating with respect to the spin rotator of the vehicle and the reactor wheel speed is maintained constant. The combination of reaction wheel and spin rotator to obtain control of several degrees of freedom also results in a complex control system which does not have the capability of compensating for vehicle or spacecraft nutation motion either created internally by operation of the system or due to outside influences, in addition to being unreliable and not suitable for missions of long duration.
In conventional prior art systems, the effect of the minimum impulse bit available from thrustors is to induce a limit cycle in the spacecraft nutation. To hold the nutation-caused attitude (pointing error) within bounds of roll and yaw limits requires frequent jet firings. For example jet firing would be required about every 240 seconds to hold the pointing error within allowable deadband tolerances. While the fuel penalty for such frequent jet firings is small, the fact that such frequent firings are required from a cold start is very detrimental to thrustor reliability if, for example, monopropellant hydrazine thrustors are employed.
One system employing jet thrustors and momentum wheels in current development attempts to resolve some of the problems stated in the preceding paragraph. It attempts to accomplish this by reducing the wheel momentum to as low a value as possible so as to make the nutation period as long as possible. This minimizes the frequency of jet firings. Such system also uses a very low thrust jet device. Reduction in the minimum impulse though accomplished, is acquired at the expense of an increase in the nutation period. The thrustors employed are hydrazine dissociation types in which the fuel in cold gas form is stored, such thrustors being effectively of the cold-gas type. The specific impulse for this type of system, as well as the thrust level, is very low in the order of 80 millipounds. As a result, to effect adequate correction or reorientation of the vehicle, another propulsion system added to the system above described is needed to obtain the higher thrust levels required.
A number of theoretical papers relating to equations of motion and generally touching on three-axes control for synchronous-orbit type communication satellites have been recently written, the most important of which is entitled "Attitude Stabilization of Synchronous Satellites Employing Narrow-Beam Antennas" by Dougherty, Lebsock and Rodden, AIAA paper number 70-457, Third Communications Satellite Conference at Los Angeles, California, April 6-8, 1970. The purely theoretical treatment of equations of motion developed in this paper is thought to be necessary to tie in with the instant specification. Consequently, the equations therein were modified, to the extent required by exigencies of the instant application, in the section of this specification entitled, "Theoretical Development and Equations of Motion," hereinbelow.
Insofar as all prior art is concerned, it may be stated in summation that no known three-axes control systems have been developed capable of being operated at synchronous altitudes, the prior art systems generally being based on dual spin technology for providing attitude control. Though three-axes control systems dependent upon magnetic torquers have been used to provide control moment, these are not suitable for the system used herein for reasons already stated.